Method and apparatus for reaction propulsion

ABSTRACT

6. A method for developing propulsive reaction at air speeds in excess of sonic velocity comprising: directing an air flow into a confined channel; mixing fuel with said air flow ; exciting successive oblique shock waves in said channel to decrease the Mach number and increase the static pressure and static temperature of the flow therein; removing boundary layer air from the channel walls adjacent the terminal ends of certain of the oblique shock waves; and exciting a shock wave substantially normal to the axis of said flow at a fixed position in said channel to induce detonation of said fuel-air mixture.

United States Patent [191 Norman et al.

1 1 Dec. 11,1973

[ METHOD AND APPARATUS FOR REACTION PROPULSION [75] Inventors: Leslie W.Norman, Scottsdale;

Skillman C. Hunter, Phoenix, both [21] Appl. No.: 88,149

[52] US. Cl 60/204, 60/270, 60/244 [51]. Int. Cl. F02k 7/10, F02k 3/02[58] Field of Search 60/356, 35.6 R]

[56] References Cited UNITED STATES PATENTS 2,911,787 11/1959 Barry o60/356 2,990,142 6/1961 Ferri 1 60/356 2,763,426 9/1956 Erwin.....60/356 2,971,330 2/1961 Clark 4 4 r 1 60/35.6 3,023,571 3/1962Pietrangeli et aim 60/356 3,040,516 6/1962 Brees 60/356 3,054,255 9/1962Stratford 60/35,6

OTHER PUBLICATIONS Recent Advances in Ramjet Combustion by Gordon L.Dugger, ARS Journal, Nov. 1959, pp. 819-827.

Primary Examiner-Robert F. Stahl Attorney-John H. G. Wallace andHerschel C.

Omohundro EXEMPLARY CLAIM 6. A method for developing propulsive reactionat air speeds in excess of sonic velocity comprising: directing an airflow into a confined channel; mixing fuel with said air flow excitingsuccessive oblique shock waves in said channel to decrease the Machnumber and increase the static pressure and static temperature of theflow therein; removing boundary layer air from the channel wallsadjacent the terminal ends of certain of the oblique shock waves; andexciting a shock wave substantially normal to the axis of said flow at afixed position in said channel to induce detonation of said fuel-airmixture.

7 Claims, 4 Drawing Figures PAIENTEUDEBI 1 m;

SHEETlQfZ INVENTORSJ SKILLMA/V C. HUNTER LESLIE W. NORMAN,

AGE/VT.

PATENTEDBHI! 1 ms SNfH Z I! 2 an! HHHHHH 1N VENTOR5: SKILL MAN 0.HUNTER, BY LESLIE W. NORMA/V,

mm Mm METHOD AND APPARATUS FOR REACTION PROPULSION The present inventionrelates to a reaction propulsion method and apparatus for operationunder condi tions of relative airflow exceeding the local propagationalspeed of sound and, more particularly, to a propulsive aerothermodynamicduct wherein combustion of a fuel-air mixture is established andmaintained by novel aerodynamic means.

In propulsive aerothermodynamic ducts as known in the prior art, ambientair is induced into the inlet portion of a duct, preferably at arelative flow velocity exceeding the local propagational speed of sound,and subsequently decelerated through a normal shock wave to subsonicvelocity in a diffuser of appropriate internal contour. (In the interestof brevity, speeds relative to the local propagational speed of sound orso-called sonic velocity", will hereinafter be identified according tothe conventional notation of the art in terms of Mach number, M We,wherein v is the flow velocity under consideration and c is the localpropagational speed of sound at the prevailing conditions of airtemperature). The subsonic air flow is then further decelerated to aspeed compatible with the establishment and maintenance of a normalcombustion process, for example, to a velocity in the region of M 0.2 toM 0.5, and fuel, introduced thereto by means of suitable injectionnozzles, is ignited by appropriate means such as an electric sparkdevice, the resulting flame being stabilized and maintained in a fixedposition relative to the duct by means of a flameholder generally takingthe form of a grid or similar structure mounted transversely to the ductaxis. The gaseous combustion product is then expanded through aconvergent-divergent nozzle so as to produce a net forward thrust and isexpelled to the ambient atmosphere via an exhaust orifice. With theexception of the so-called pulse jet, which depends for its operation ona cyclically operating valve cooperative with a duct tuned to thenatural frequency thereof, aerothermodynamic ducts of the prior art,generally known as ram jets, conform broadly to the described operatingcycle, numerous refinements and variations in the design details thereofhaving been developed by those skilled in the art.

The present invention comprehends a novel aerothermodynamic ductutilizing a mode of fuel oxidation differing from the normal combustionprocess and operating independently of auxiliary flameholding orignition structures as taught in the prior art. Briefly, the inventiondepends for its operation on the combustion process generally referredto as detonation; that is, a process wherein the flame front isestablished and maintained by the temperature rise occurring across ashock wave, rather than being advanced by heat conduction as is the casein conventional internal combustion processes such as occur in ram jetas well as turbojet engines and in reciprocating engines operating onthe Otto cycle. Whereas detonation as normally considered in connectionwith an Otto cycle engine, for example, is a spacially mobile andtemporally transient phenomenon of extremely short duration, in thepresent invention a continuous, steady state detonation is establishedand maintained across a standing shock wave of fixed position relativeto a confining structure therefor, subsequent expansion of the gaseousdetonation product being used to develop a continuous thrust.

It is an object of the present invention, therefore, to provide anaerothermodynamic duct wherein propulsive force or thrust is developedby continuous detonation of a fuel-air mixture.

It is another object of the invention to provide an aerothermodynamicduct wherein combustion of a fuel-air mixture takes place across a shockwave of substantially fixed position relative to the adjacent structure.

It is another object of the invention to provide an aerothermodynamicduct wherein combustion of a fuel-air mixture takes place in consequenceof the temperature rise across a shock wave.

It is another object of the invention to provide an aerothermodynamicduct wherein combustion of a fuel-air mixture is initiated byaerodynamic means independently of auxiliary igniter means.

It is another object of the invention to provide an aerothermodynamicduct wherein combustion of a fuel-air mixture is maintained byaerodynamic means independently of an auxiliary flame holding structure.

It is another object of the invention to provide an aerothermodynamicduct wherein fuel may be introduced into the air stream substantiallyupsteam from the locus of combustion.

It is another object of the invention to provide an aerothermodynamicduct for continuous detonation of a mixture of air and gaseous fuel.

It is another object of the invention to provide an aerothermodynamicduct for continuous detonation of a mixture of air and fuel, whereinsaid fuel is hydrogen.

It is another object of the invention to provide an aerothermodynamicduct incorporating means for the excitation of relatively stationaryshock waves internally thereof, thereby to provide sufficienttemperature rise to initiate and maintain continuous detonation of afuel-air mixture.

It is another object of the invention to provide an aerothermodynamicduct wherein continuous detonation of a fuel-air mixture may beaccomplished by means of a system of interrelated shock waves that maybe partially external to the confining walls of the duct.

Still further objects of the invention will be made apparent in thefollowing portions of this specification which, taken in conjunctionwith the appended drawings, are particularly descriptive of an exemplaryembodiment thereof.

In the drawings, which are merely illustrative and not to be construedby way of limitation, and in which like elements are denoted by likereference numerals,

FIG. 1 is a side elevation, partially broken away, of an aircraftincorporating a propulsive engine according to the present invention;

FIG. 2 is a bottom plan view of the aircraft shown in FIG. 1;

FIG. 3 is a sectional plan view on an enlarged scale taken along theline 3-3 of FIG. 1 showing certain internal features of the propulsiveengine; and

FIG. 4 is a sectional elevation on the same scale as FIG. 3 and takenalong the line 44 of FIG. 2.

In the drawing, an aircraft 1 of a type adapted for flight at so-calledhypersonic Mach numbers, for example, in the region of M 6, is shown ashaving an acutely tapered nose portion 2, an aerodynamic sustainingsurface or wing 3 having sharply swept leading edges 4 and a body orfuselage 5, the lower portion of the latter housing a propulsive engine10, according to the present invention, it being understood, however,that the general configuration of the aircraft 1 forms no part of theinvention except insofar as it is adapted to co-operate with engine in amanner to be more particularly described hereinafter.

As is more clearly shown in FIGS. 3 and 4, the engine 10 comprises anelongated duct having an intake portion 11, an intermediate portion 12adapted to excite and contain a system of shock waves in a manner to bedescribed hereinafter, a combustion portion 13 including a hingedlyadjustable ramp 14, and an exhaust nozzle 15. Associated with the engine10 is a second engine 20, which may, for example, be a turbo-jet ofknown type, the purpose and operation of which will be made apparenthereinafter.

As previously discussed, the propulsive engine 10, which is anaerothermodynamic duct, depends for its operation on the continuousdetonation of a fuel-air mixture due to the temperature rise across astanding shock wave, the latter preferably being a normal shock wavethough not necessarily limited thereto, the shock wave across which thedetonation takes place in the embodiment shown in the drawings beingindicated by the dashed line 62. The structural elements and aerodynamicphenomena leading to the development and determining the location ofthis shock wave will now be described with particular reference to thecombination of vehicle and motor illustrated, the aerodynamic phenomenabeing described for exemplary purposes only with reference to a freestream Mach number of 6.5 at an altitude of 125,000 feet, it beingunderstood, however, that where steady state operation under differentambient conditions is desired appropriate modifications to particularfeatures of the design may be made in accordance with aerodynamicprinciples well known to those skilled in the art.

Referring to FIG. 1, the dashed line 21 indicates an oblique shock waveoriginating at the nose 2 and extending rearwardly therefrom at an angledetermined by the free stream Mach number, the intake portion 11 of themotor 10 being provided with a forwardly raked induction scoop 22 havinga leading edge 23 positioned so as to intersect the shock wave 21 asshown. As is well known to those skilled in the art, the Mach numbercharacterizing the flow in the region designated by the numeral 24behind the shock wave 21 will be somewhat lower than the free streamvalue of 6.5 previously stated, this decrease in Mach number beingaccompanied by a rise in static pressure and a similar rise in statictemperature; for convenient reference throughout the followingdiscussion, the values of pressure, temperature, Mach number andvelocity prevailing in the various flow regions shown in the drawing aretabulated below:

TABLE 1 Total Statlc temp. temp. Mach Velocity R.) C R.) No. (M)(t.p.s.)

3, 850 460 8. 50 6, B 3, 850 542 5. 88 6, B70 3, 850 614 5. 52 7, 2 3,850 975 4. 18 6, 450 3, 850 1, 208 3. 64 6, 200 3, B50 1, 437 3. 24 6,030 3, 650 1, 820 2. 5O 5, 400 5, 385 4, 950 0. 801 2, 760 5, 385 1,0705. 0 8, 335

In accordance with the reduced Mach number prevailing in region 24, theaircraft fuselage S is provided with a ramp 25 having a sharp leadingedge positioned transversely of the relative flow so as to generate asec- 5 0nd oblique shock wave 26 extending rearwardly therefrom tointersect with shock wave 21 at the leading edge 23 of induction scoop22, the ramp 25 preferably being provided with boundary layer diversionslots 27 of known type in order to minimize the deflection of turbulentflow into the region 28 behind shock wave 26. Referring to Table I, itwill be seen that the Mach number prevailing in region 28 has been stillfurther reduced and that the static pressure, and static temperaturehave been correspondingly increased, the total temperature remainingunchanged since the aerodynamic process characterizing the shock wave 26is adiabatic.

In consequence of the deflection of the air flow at leading edge 23, athird oblique shock wave 29 extends rearwardly therefrom into intakeporton 11, the pressure, temperature and Mach number prevailing inregion 30 behind shock wave 29 being further changed as indicated inTable I. As shown in FIG. 4, shock wave 29 intersects the entrance tointermediate duct portion 12 adjacent the leading edge 31 of a secondboundary layer diversion structure comprising wall members 32. and 33,34 and 35, and 36 and 37 defining diversionary passages 38, 39 and 40,respectively, the last named preferably being of bifurcated constructionso as to merge with passages 38 and 39, as through symmetricallydisposed transition passages 41. The intermediate duct portion 12extends rearwardly with substantially constant vertical dimension buttapers in the horizontal plane as shown in FIG. 3, thereby tending toestablish and maintain a symmetrically disposed pair of oblique shockwaves 42 defining flow regions 43 and intersecting with each other inmutually reinforcing manner so as to define another flow region 44characterized by temperatures, pressures and Mach number as shown inTable I. As the flow in region 44 encounters the discontinuity of wallcontour occurring at the leading edge of hingedly mounted ramp 14 itgives rise to an oblique shock wave 52 extending rearwardly therefromthrough combustion portion 13 so as to intersect upper wall member 54defining boundary layer diversion passage 55, the pressure, temperatureand Mach number prevailing in flow region 53 behind shock wave 52 beingas indicated in Table I.

I-Iingedly mounted ramp 14 defining the bottom wall of combustionportion 13 is angularly adjustable by means of external controlmechanism 60 co-operative with linkage 61 to define a throat in theinlet region of converging-diverging nozzle 80 of appropriate dimensionsfor maintaining the normal shock wave 62 across which detonation of thefuel-air mixture takes place, the resulting flow conditions in region 63immediately therebehind being as indicated in Table l.

The formation of shock wave 62, is initiated by struc ture remote fromnozzle 80, as will now be described. Referring back to intake portion11, the bottom wall 70 thereof includes a hingedly mounted door 71normally lying flush therewith as shown, but angularly displaceable bymeans of a suitable control actuator for projection into the inducedairflow as indicated by dashed lines 71', the leading edge of the door71 being adapted to give rise to a normal shock wave, as indicated bydashed line 62' when in the displaced position, and to permit thedetachment of said shock wave when moved to the retracted or flushposition. Thus, when the aircraft 1 has been accelerated to the desiredMach number at the desired altitude, as for example by means of ajettisonable booster rocket not shown, the door 71 may be momentarilydisplaced by means of a suitable external control actuator therefor tothe position 71, thereby to promote the formation of normal shock wave62, the door 71 being thereupon returned to its flush position so as topermit the detachment therefrom of the shock wave 62; thus the door 71operates as a starting device only, taking no part in the steady stateoperation of the engine. The detached shock wave 62' is carried by themotion of the induced air so as to move downstream through the ductuntil it reaches the position 62 where it is retained by the step formedat the trailing edge of hinged ramp l4 and the inlet region of nozzle80, as hereinbefore described.

Referring to Table I, it will be seen that a stepwise increase in statictemperature accompanies each successive shock wave in the aircraft andduct system. To take advantage of this increase in the mannercomprehended by the present invention, a fuel having a detonationtemperature below the static temperature occurring in the normal shockwave 62 must be selected, an exemplary fuel meeting this condition forthe present embodiment of the invention being hydrogen. As shown inTable l, the static temperature occurring in flow region 53 downstreamof oblique shock wave 52 may be higher than the temperature required forinitiation of combustion in the air-hydrogen mixture, the time requiredfor the mixture to traverse this region being so short, however, incomparison with the time needed to achieve combustion, that the mixturemay remain chemically unaffected until it reaches the nor mal shock wave62, wherein the static temperature rises abruptly to a valueapproximately equal to the preceding value of total temperature withconsequent detonation of the mixture and a further static temperatureincrease to 4950R as shown in Table I. If the operating conditions tendto deviate from those for which the duct and its associated shock wavesystem are designed, pre-ignition may occur in region 53, and in suchevent the angular position of hinged ramp 14 may be adjusted so as toincrease the Mach number prevailing therein, such an increase in Machnumber operating to reduce the static temperature and thus restore thecombustion process to the region of the normal shock wave 62 whereindetonation is excited.

lnasm uch as no static temperature equal to or greater than the ignitiontemperature of hydrogen exists upstream of the shock wave 62 in thesteady state operation of the duct system, it will be apparent that thefuel may be introduced into the airflow at any point forward of suchsteady state position. Accordingly, in the preferred practice of theinvention, fuel is introduced into the air flow substantially upstreamfrom the combustion portion 13, thereby to take advantage of the mixingafforded by the system of successive shock waves excited by the severalramps and area changes of the duct system, a suitable arrangement ofinjection nozzles comprising, for example, a plurality of spaced,vertically aligned members 75 located as shown at the inlet region ofintermediate duct portion 11, the members 75 being of suitablecross-sectional shape, for example, symmetrically rhombic, and havingdischarge orifices provided in the trailing edges thereof. As will beapparent to those skilled in the art, however, numerous other locationsfor such nozzles may be provided, the operating principles andaerodynamic flow patterns characteristic of the invention making itfeasible even to inject fuel upstream of the induction scoop 22 if sodesired.

The boundary layer diversion passages, 38, 39, 50 and 55 provided bywall members 33, 34, 51 and 54 serve to mitigate the formation, adjacentthe duct walls, of unduly thick boundary layers, such as might tend tochoke the airflow, distort the desired shock wave pattern and reduce theoperating efficiency of the engine, the induced air, diverted bypassages 38 and 39 being directed to auxiliary convergent-divergentnozzles 78 and 79 where it is vented to the exhaust nozzle 15 and mergeswith the primary flow expanding thereinto through convergent-divergentnozzle 80 provided at the rear of combustion portion 12, the internalcontour of the nozzle 15 of the instant embodiment of the in ventionbeing formed as shown, sufficient change in flow momentum to developsatisfactory thrust being imparted by virtue of the high velocity of theexhaust gas as shown in Table l.

Though the vehicle shown may be accelerated to the desired operatingMach number by auxiliary means, such as a jettisonable booster, ashereinbefore stated, it will be readily apparent that landing cannot beaccomplished safely and satisfactorily at a velocity corresponding tothe high Mach number required for steady state operation. For landingthe vehicle, therefore, second engine 20 is provided, the intake passage82 and exhaust passage 83 thereof being integrated with the intermediateportion 12 and exhaust nozzle 15 of the primary engine 10 andconnectable in flow communication therewith by means of appropriatelyhinged doors 84 and 85 operable by means of suitable control actuatorsof known type, not shown, to open positions as indicated by broken lines84' and 85' respectively. Thus, when it is desired to land the vehicle,the combustion process in the primary engine 10 may be suspended byinterruption of the fuel supply thereto, thereby permitting the vehicleto decelerate under the influence of aerodynamic drag until a speed isattained at which it is safe to open doors 84 and 85 so as to divert aportion of the ducted air flow to the turbo-jet engine 20 to permitstarting and operation thereof according to the conventional proceduresfor such propulsion units, the landing operation then being carried outunder the power of engine 20 at such speed as the aerodynamiccharacteristics of vehicle 1 may require.

it will be apparent to those skilled in the art that while thepropulsion method and apparatus of the present invention have beendescribed with reference to operation in a specific airframe, at aspecific Mach number and altitude, and with a specific fuel, and withair or ambient atmosphere as a specific oxidant, the utility of theinvention is not limited thereto but is adaptable, according to knownprinciples of aeroand thermo-dynamics, to operation in conjunction witha variety of vehicles or other co-operating structure, under acorresponding variety of superand hypersonic flow conditions, and with avariety of known fuels and atmospheres such as methane and ammonia; inthe latter case, of course, the aerodynamic fluid is the fuel, and theoxidant is injected into it for combustion. It is, therefore,anticipated that those skilled in the art will have occasion to practicenumerous variations on specific features of the propulsive method andapparatus disclosed herein, the instant description and drawings beingpresented by way of example only and not by way of limitation, and it isour desire that all such variations falling within the spirit and scopeof the invention be secured to us by Letters Patent.

We claim:

1. In combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: a duct for confining aflow of air, said duct including a forwardly directed inlet portion anda rear wardly directed exhaust portion; means for mixing fuel with saidflow of air; constrictive means intermediate said inlet portion and saidexhaust portion for retaining a shock wave having a temperature risethereacross sufficient to cause spontaneous detonation of the mixture ofsaid air and said fuel; and means spaced longitudinally of the vehicleand said duct and operative to create successive oblique shock waves,part of said means being disposed to cause the oblique wave createdthereby to engage the duct wall at one side of said constrictive meansand being adjustable to vary the effective area of the duct at saidrestrictive means.

2. In combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: a duct for confining aflow of air, said duct including a forwardly directed inlet and arearwardly directed outlet; nozzle means positioned in said ductdownstream of said inlet, said nozzle means being operable to introducefuel into said flow of air for mixture therewith; and constrictive meanspositioned in said duct downstream of said nozzle means for retaining ashock wave having a temperature rise thereacross sufficient to causespontaneous detonation of the mixture of said air and said fuel, andmeans for generating a detonation-causing shock wave, said meansincluding a ramp to create an oblique shock wave, said ramp beinghingedly adjustable about a line of oblique shock wave initiation tovary the effective cross-sectional area of said duct at saidconstriction where said detonationcausing shock wave is retained.

3.1a combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: a duct for confining aflow of air, said duct including a forwardly directed inlet and arearwardly clirected outlet; means in said duct downstream of said inletfor mixing fuel with said air; a constricted throat downstream of saidmixing means for creating and retaining a normal detonation-causingshock wave in said duct; and means for creating an oblique shock waveupstream of said throat and varying the effective size of said throat,said means providing a ramp disposed to cause an oblique wave to engagethe duct wall at one side of said restricted throat in all effectivesizes of the latter.

4. in combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: means defining an airduct adjoining said vehicle, said duct including a forwardly directedinlet and a rearwardly directed outlet door means disposed in a wall ofsaid duct downstream of said inlet, said door means being hingedlyoperable from a closed position to an open position to provide anopening in said wall and to constrict the flow area internally of saidduct for exciting a normal shock wave therein, and from said openposition to said closed position to permit said wave to be propagateddownstream of said door; means for introducing fuel into said ductdownstream of said inlet for mixture with said air; and a constrictionin said duct downstream of said injection means for arresting saidnormal shock wave and retaining the same in fixed position relative tosaid duct, thereby to excite detonation of said fuel air mixture in aregion upstream of said outlet.

5. In combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine com prising: means defining an airduct adjoining said vehicle, said duct including a forwardly directedinlet and a rearwardly directed outlet; means upstream of said inlet forexciting a first shock wave positioned for cooperation with said inlet;door means disposed in a wall of said duct downstream of said inlet,said door means being hingedly operable from a closed position to anopen position to provide an opening in said wall and to constrict theflow area internally of said duct for exciting a normal shock wavetherein, and from said open position to said closed position to permitsaid wave to be propagated downstream of said door; means forintroducing fuel into said duct downstream of said inlet for mixturewith said air; and a constriction in said duct downstream of saidinjection means for arresting said normal shock wave and retaining thesame in fixed position relative to said duct, thereby to excitedetonation of said fuel air mixture in a region upstream of said outlet.

6. A method for developing propulsive reaction at air speeds in excessof sonic velocity comprising: directing an air flow into a confinedchannel; mixing fuel with said air flow; exciting successive obliqueshock waves in said channel to decrease the Mach number and increase thestatic pressure and static temperature of the flow therein; removingboundary layer air fron the channel walls adjacent the terminal ends ofcertain of the oblique shock waves; and exciting a shock wavesubstantially normal to the axis of said flow at a fixed position insaid channel to induce detonation of said fuel-air mixture.

7. in combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: a duct for confining aflow of air, said duct including a forwardly directed inlet and arearwardly directed outlet; means in said duct downstream of said inletfor mixing fuel with said air; means spaced longitudinally of said ductfor creating successive oblique shock waves to decrease the Mach numberand increase the static pressure and static temperature of the flow inthe duct; means disposed in said duct to that required for spontaneousdetonation for removing boundary layer air from the duct walls adjacentpoints of impact of certain oblique shock waves therewith; a constrictedthroat in said duct downstream of said mixing means for retaining anormal shock wave across which a temperature rise sufficient to causespontaneous detonation of the mixture of air and fuel occurs; andadjustable means for varying the effective area of said constrictedthroat.

UNITED STATES PATENT OFFICE CERTIFICATE OF CORRECTION Dated December 11,1973 Patent No. 3,777,487

Inventor(s) Leslie Norman and Skillrnan C. Hunter It is certified thaterror appears in the above-identified patent and that said LettersPatent are hereby corrected as shown below:

Claim 4, line 5, after "outlet" insert -semicolon Claim 6, line'7,change "iron" to -from--.

Claim 7, lines 10 and 11, after "duct" (second occurrence) delete tothat required for spontaneous detonation";

line 10, after "duct" (first occurrence) insert -'to that required forspontaneous detonation--.

Signed and sealed this 21st day of May 197M.

(SEAL) Attest EDWARD, IE.EI.ETCZ5R,JR. C. i-IAHSHALL DAN'N AttestingOfficer Commissioner of Patents uscoMM-oc man-ps9 9 U. 5. GOVERNMENYPRINTING OFFICE III O-J'C-Jl,

F ORM PO-1 050 (10-69) UNITED STATES PATENT OFFICE CERTIFICATE OFCORRECTION Patent No. 3 ,777,487 Dated December 11, 1973 Inventor(s)Leslie Norman and Skiilman C. Hunter It is certified that error appearsin the above-identified patent and that said Letters Patent are herebycorrected as shown below:

Claim 4, line 5, after "outlet" insert -sernico1on Claim 6, line 7,change "iron" to -from---.

I Claim 7, lines 10 and 11, after "duct" (second occurrence) delete "tothat required forspontaneous detonation";

line 10, after "duct" (first occurrence) insert --to that required forspontaneous detonation--.

Signed and sealed this 21st day of May 1971 (suAL) Attest Bill/AND ii.L'LE TGIil-JR JR C MARSHALL BARN Attesting Ufficer Commissioner ofPatents RM F'0-1050 (10-69) USCOMM-DC 60376-P69 9 U S. GOVEHNMENYPRINTING OFFICE 909 O386-3!l.

1. In combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: a duct for confining aflow of air, said duct including a forwardly directed inlet portion anda rearwardly directed exhaust portion; means for mixing fuel with saidflow of air; constrictive means intermediate said inlet portion and saidexhaust portion for retaining a shock wave having a temperature risethereacross sufficient to cause spontaneous detonation of the mixture ofsaid air and said fuel; and means spaced longitudinally of the vehicleand said duct and operative to create successive oblique shock waves,part of said means being disposed to cause the oblique wave createdthereby to engage the duct wall at one side of said constrictive meansand being adjustable to vary the effective area of the duct at saidrestrictive means.
 1. In combination with a vehicle for travel throughair at supersonic speed, a reaction propulsion engine comprising: a ductfor confining a flow of air, said duct including a forwardly directedinlet portion and a rearwardly directed exhaust portion; means formixing fuel with said flow of air; constrictive means intermediate saidinlet portion and said exhaust portion for retaining a shock wave havinga temperature rise thereacross sufficient to cause spontaneousdetonation of the mixture of said air and said fuel; and means spacedlongitudinally of the vehicle and said duct and operative to createsuccessive oblique shock waves, part of said means being disposed tocause the oblique wave created thereby to engage the duct wall at oneside of said constrictive means and being adjustable to vary theeffective area of the duct at said restrictive means.
 2. In combinationwith a vehicle for travel through air at supersonic speed, a reactionpropulsion engine comprising: a duct for confining a flow of air, saidduct including a forwardly directed inlet and a rearwardly directedoutlet; nozzle means positioned in said duct downstream of said inlet,said nozzle means being operable to introduce fuel into said flow of airfor mixture therewith; and constrictive means positioned in said ductdownstream of said nozzle means for retaining a shock wave having atemperature rise thereacross sufficient to cause spontaneous detonationof the mixture of said air and said fuel, and means for generating adetonation-causing shock wave, said means including a ramp to create anoblique shock wave, said ramp being hingedly adjustable about a line ofoblique shock wave initiation to vary the effective cross-sectional areaof said duct at said constriction where said detonation-causing shockwave is retained.
 3. In combination with a vehicle for travel throughair at supersonic speed, a reaction propulsion engine comprising: a ductfor confining a flow of air, said duct including a forwardly directedinlet and a rearwaRdly directed outlet; means in said duct downstream ofsaid inlet for mixing fuel with said air; a constricted throatdownstream of said mixing means for creating and retaining a normaldetonation-causing shock wave in said duct; and means for creating anoblique shock wave upstream of said throat and varying the effectivesize of said throat, said means providing a ramp disposed to cause anoblique wave to engage the duct wall at one side of said restrictedthroat in all effective sizes of the latter.
 4. In combination with avehicle for travel through air at supersonic speed, a reactionpropulsion engine comprising: means defining an air duct adjoining saidvehicle, said duct including a forwardly directed inlet and a rearwardlydirected outlet door means disposed in a wall of said duct downstream ofsaid inlet, said door means being hingedly operable from a closedposition to an open position to provide an opening in said wall and toconstrict the flow area internally of said duct for exciting a normalshock wave therein, and from said open position to said closed positionto permit said wave to be propagated downstream of said door; means forintroducing fuel into said duct downstream of said inlet for mixturewith said air; and a constriction in said duct downstream of saidinjection means for arresting said normal shock wave and retaining thesame in fixed position relative to said duct, thereby to excitedetonation of said fuel air mixture in a region upstream of said outlet.5. In combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: means defining an airduct adjoining said vehicle, said duct including a forwardly directedinlet and a rearwardly directed outlet; means upstream of said inlet forexciting a first shock wave positioned for cooperation with said inlet;door means disposed in a wall of said duct downstream of said inlet,said door means being hingedly operable from a closed position to anopen position to provide an opening in said wall and to constrict theflow area internally of said duct for exciting a normal shock wavetherein, and from said open position to said closed position to permitsaid wave to be propagated downstream of said door; means forintroducing fuel into said duct downstream of said inlet for mixturewith said air; and a constriction in said duct downstream of saidinjection means for arresting said normal shock wave and retaining thesame in fixed position relative to said duct, thereby to excitedetonation of said fuel air mixture in a region upstream of said outlet.7. In combination with a vehicle for travel through air at supersonicspeed, a reaction propulsion engine comprising: a duct for confining aflow of air, said duct including a forwardly directed inlet and arearwardly directed outlet; means in said duct downstream of said inletfor mixing fuel with said air; means spaced longitudinally of said ductfor creating successive oblique shock waves to decrease the Mach numberand increase the static pressure and static temperature of the flow inthe duct; means disposed in said duct to that required for spontaneousdetonation for removing boundary layer air from the duct walls adjacentpoints of impact of certain oblique shock waves therewith; a constrictedthroat in said duct downstream of said mixing means for retaining anormal shock wave across whiCh a temperature rise sufficient to causespontaneous detonation of the mixture of air and fuel occurs; andadjustable means for varying the effective area of said constrictedthroat.